The control of an aircraft generally calls on many control systems, including in particular:                the primary flight control system, which makes it possible to control the aircraft's movement around its roll, yaw and pitch axes, by acting on the ailerons, rudders, elevators and the trimmable horizontal stabilizer (THS). It also controls the aircraft's drag by acting on the spoilers;        the secondary flight control system, making it possible to control the camber of the wings and therefore the lift, during takeoff and landing phases, by acting on the flaps and the slats, also called leading edge slats;        the propulsion control system, making it possible to control and reverse the thrust from the engines;        the brake control system of the landing gear;        the steering control system;        the control system for the hydraulic circuits.        
In general the primary flight system is simply called the flight control system. We will adopt this convention below in order to simplify the description.
The flight control system connects the steering members (control column, rudder bar, etc.) and the aerodynamic tip-control surfaces (ailerons, vertical stabilizers, elevators, etc.). Modern jetliners have electric-type flight control systems in which the mechanical actions on the piloting members are converted into analog signals that are sent to actuators maneuvering the control surfaces.
FIG. 1 diagrammatically illustrates the architecture of a flight control system 100, known from the state of the art. We have shown a piloting member 110, for example a side-stick, equipped with one or more sensors 115, for example position sensors and/or angular sensors providing position and/or orientation information to the flight control computer 120. The computer 120 determines, from information received from the various piloting members 110, here including the auto-pilot (not shown) and/or, if applicable, airplane sensors 150 (accelerometer, rate gyro, inertial unit), the flight controls to be applied to the actuators 130. These actuators are typically hydraulic cylinders controlled by servo-valves or electric motors acting on the aerodynamic flight-control surfaces of the aircraft 140. The actuators 130 on the one hand, and the aerodynamic flight-control surfaces 140 on the other hand, are equipped with sensors respectively denoted 135 and 145. These sensors inform the computer 120 on the positions and/or orientations of the mobile elements of the actuators as well as those of the control surfaces. For example, one sensor 135 could indicate the translational position of a cylinder, one sensor 145 the orientation of a flap.
The computer 120 has both a command function and a monitoring function. It is connected to the actuators by first cables 133 intended to transmit the analog control signals. It is also connected to the sensors 135 and 145 respectively equipping the actuators and the control surfaces themselves by second cables 137 and third cables 147. It can thus, at any time, monitor the status of the actuators and verify that the commands have been carried out correctly.
In reality, a flight control system is generally made up of several independent computers, each computer having its own set of sensors and actuators and its own cable network.
FIG. 2 diagrammatically illustrates the architecture of an aircraft control system 200. It comprises the flight control system as well as a plurality of other control systems, examples of which were provided in the introduction. For simplification purposes, only two control systems SC1 and SC2 have been shown here.
Each control system SCn, n=1,2 comprises at least one dedicated computer 210 n processing the signals received from one or more sensor(s) 220n, and transmitting commands to one or more actuator(s) 230 n via a plurality of cables.
The different control systems are placed in the avionics bay (delimited in the diagram by a broken double line) and connected to each other using an avionics network, for example an AFDX (Avionics Full DupleX switched Ethernet) network. It will be recalled that the AFDX network, specifically developed for aeronautics needs, is based on a switched Ethernet network. A detailed description of the characteristics of this network can be found in a document entitled “AFDX protocol tutorial” and in patent application FR 2001-0014263 (published as FR-A-2832011 on May 9, 2003)filed Nov 5, 2001 in the Applicants name.
When the computer of the control system SC2 needs a measurement on a piece of equipment E, done by a sensor belonging to the control system SC1, the computer 2101 in charge of SC1 transmits that measurement to computer 2102 through the avionics network. However, it is sometimes necessary, to satisfy availability constraints (e.g. breakdown of computer 2101) or latency constraints (transfer time via computer 2101) to duplicate the sensor on the equipment E. In other words, a second sensor 2202 is then provided on the equipment E, directly connected to the computer SC2.
This aircraft control system architecture has a certain number of drawbacks, including the multiplication of the number of sensors and related cables, which strains the aircraft's weight budget. Furthermore, the analog signals transmitted by the sensors can be affected by noise due to electromagnetic disturbances.
The object of the present invention is therefore to propose an aircraft control system resolving the aforementioned drawbacks, i.e. a system making it possible to reduce the cabling between the avionics bay and the sensors/actuators, as well as to reduce the number of sensors on the equipment, without sacrificing the required level of safety.